Evaluation of Electronics Shielding in
Micro-satellites
L. Varga and E. Horvath
Defence R&D Canada - Ottawa
TECHNICAL MEMORANDUM
DRDC Ottawa TM 2003-017
February 2003
Evaluation of Electronics Shielding in
Micro-satellites
L. Varga
DRDC Ottawa
E. Horvath
JERA Consulting
Defence R&D Canada - Ottawa
Technical Memorandum
DRDC Ottawa TM 2003-017
February 2003
© Her Majesty the Queen as represented by the Minister of National Defence, 2003
© Sa majesté la reine, représentée par le ministre de la Défense nationale, 2003
Abstract
This report investigates radiation shielding capabilities of micro-satellite bus model structures,
incorporating different designs and materials for protecting internal spacecraft electronics
from the ionizing radiation of the space environment. The modelling calculations have been
carried out with a 3D Monte Carlo radiation transport code. The results indicate that the
greatest reduction of total ionizing dose (TID) is observed with traditional aluminum
spacecraft structures, although structures made with lighter poly-carbon materials with added
thin layer of high-Z material can provide comparable radiation protection in addition to some
spacecraft mass reduction.
Résumé
Ce rapport étudie les possibilités d'armature de rayonnement des structures de modèle de
micro-satellite, de différentes conceptions d'incorporation et des matériaux pour protéger
l'électronique interne de vaisseau spatial contre la radiation ionisante de l'environnement de
l'espace. Les calculs modelants ont été effectués avec le code de transport derayonnement de
3D Monte Carlo. Les résultats indiquent que la plus grande réduction de la dose s'ionisante
totale (TID) est observée avec les structures traditionnelles de vaisseau spatial d'aluminum,
bien que, les structures faites avec des matériaux plus légers de poly-carbone avec la couche
mince supplémantaire du haut-Z matériel puissent assurer la radioprotection comparable en
plus d'une certaine réduction de la masse de vaisseau spatial.
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Executive summary
Electronic devices located inside orbiting satellites in the near-Earth space environment are
exposed during the mission to ionizing space radiation propagating through the micro-satellite
structure and creating damage. Because of such damage, over a period of time electronic
components can fail thus jeopardising the mission success. A trade-off exists between the
amount of shielding the spacecraft needs for protection from the space radiation effects and
the mass reduction effort to reduce the launch cost. The results of this study indicate that new
lighter materials can be utilized to reduce the weight budget of the mission. Materials such as
poly-carbon PEEK can be used to build the micro-satellite structure. Shielding effectiveness
can be also improved by lining the interior of the spacecraft structural panels with a thin layer
of high-Z material such as tantalum. This design can provide protection that is comparable to
traditional aluminum structures but also can lead to weight reduction and thus reduction in
launch cost.
Varga L., Horvath E. 2003. Evaluation of Electronic Shielding in Micro-satellites. DRDC
Ottawa TM 2003-017. Defence R&D Canada - Ottawa .
DRDC Ottawa TM 2003-017
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Sommaire
Des satellites orbitaux intérieurs localisés par dispositifs électroniques dans l'environnement
de l'espace de la proche-Terre sont exposés pendant la mission au rayonnement s'ionisant de
l'espace propageant par la structuremicro-satelliteet créant des dommages. En raison d'un tel
dommages, sur une certaine période de temps les composants électroniques peut échouer de
ce fait compromettant le succès de mission. Une compensation existe entre la quantité de
protéger les besoins de vaisseau spatial de protection contre les effets de rayonnement de
l'espace et l'effort de masse de réduction de réduire le coût de lancement. Les résultats de cette
étude indiquent que de nouveaux matériaux d'allumeur peuvent être utilisés pour réduire le
budget de poids de la mission. Des matériaux tels que le PEEK de poly-carbone peuvent être
employés pour établir la structure satellite. L'armature de l'efficacité peut être également
améliorée en rayant l'intérieur des panneaux structuraux de vaisseau spatial avec une couche
mince de haut-Z matériel tel que le tantale. Cette conception peut assurer la protection qui est
comparable aux structures traditionnelles d'aluminum mais également peut mener à la
réduction de poids et ainsi à la réduction en coût de lancement.
Varga L., Horvath E. 2003. Evaluation of Electronic Shielding in Micro-satellites. DRDC
Ottawa TM 2003-017. R & D pour la défense Canada – Ottawa.
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Table of contents
Abstract........................................................................................................................................ i
Executive summary ................................................................................................................... iii
Sommaire................................................................................................................................... iv
Table of contents ........................................................................................................................ v
List of figures ............................................................................................................................ vi
List of Tables........................................................................................................................... viii
1. INTRODUCTION ................................................................................................................ 1
2. MICRO-SATELLITE STRUCTURE SCHEMES............................................................ 2
3. SHIELDING EFFECTIVENESS........................................................................................ 6
4. SOLAR FLARE EFFECT ................................................................................................. 12
5. ENERGY WINDOW CONTRIBUTION ......................................................................... 14
6. SHIELDING BY LOCATION .......................................................................................... 15
7. DISCUSSION...................................................................................................................... 17
8. SUMMARY......................................................................................................................... 20
9. REFERENCES ................................................................................................................... 20
List of symbols/abbreviations/acronyms .................................................................................. 21
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List of figures
Figure 1. Schematic picture of the interior of the model micro-satellite, showing the locations
of the electronic housings... ........................................................................................................ 2
Figure 2. Traditional aluminum micro-satellite bus structures schemes. Structure “B” models
“spot” shielding by increasing the electronic housing thickness... ............................................ 3
Figure 3. The solid aluminum panel is replaced with honeycomb aluminum panel in “C” and
with carbon composite PEEK honeycomb in “D”. Both honeycomb panels are covered with
0.1mm aluminum sheeting inside and outside. The electronic housings are 1mm thick solid
aluminum.....................................................................................................................................4
Figure 4. Carbon composite PEEK has replaced aluminum material in structural panels,
aluminum electronic housings, however, remain. A thin coating (0.1mm) of high Z value
material tantalum is added to the interior surface in “F”............................................................4
Figure 5. Electronic housings are attached to the structural panels. The panels have the same
composition as in structure “F”. ................................................................................................. 5
Figure 6. Model representation of the flex-board structure panel. Aluminum housings have
been eliminated..................................................................................................................................... .. 5
Figure 7. Calculated annual trapped electrons TID values into micro-satellite structures A to H
and unprotected solid-state device, trace "I", for LEO, MEO and LEO-GEO transfer orbits
with inclination angle of 30 degrees. ......................................................................................... 7
Figure 8. Calculated annual trapped electrons TID values into micro-satellite structures A to H
and unprotected solid-state device, trace "I", for LEO, MEO and LEO-GEO transfer orbits
with inclination angle of 60 degrees. ........................................................................................ 7
Figure 9. Calculated annual trapped electrons TID values into micro-satellite structures A to H
and unprotected solid-state device, trace "I", for LEO, MEO and LEO-GEO transfer orbits
with inclination angle of 60 degrees Figure Name. .................................................................... 8
Figure 10. Calculated trapped protons annual TID values into micro-satellite structures A to H
and unprotected solid-state device, trace "I", for LEO, MEO and LEO-GEO transfer orbits
with inclination angle of 30 degrees........................................................................................... 9
Figure 11. Calculated trapped protons annual TID values into micro-satellite structures A to H
and unprotected solid-state device, trace "I", for LEO, MEO and LEO-GEO transfer orbits
with inclination angle of 60 degrees. ......................................................................................... 9
Figure 12. Calculated trapped protons annual TID values into micro-satellite structures A to H
and unprotected solid-state device, trace "I", for LEO, MEO and LEO-GEO transfer orbits
with inclination angle of 85 degrees......................................................................................... 10
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Figure 13. Relative shielding effectiveness of the micro-satellite model structures A to F for
two specific orbit scenarios. The ratio is with respect to the fully exposed solid-state device..11
Figure 14. Calculated solar protons annual TID values into micro-satellite model structures A
to H and exposed solid-state device (shown for reference as trace 'I"). No contribution to TID
is observed at low altitude orbits and low angle of inclination. ............................................... 12
Figure 15. Calculated solar protons annual TID values into micro-satellite structures A to H
and unprotected solid-state device (shown for reference as trace 'I") at medium orbit
inclination angle. ...................................................................................................................... 13
Figure 16. Calculated solar proton annual TID values into micro-satellite structures A to H
and unprotected solid-state device (shown for reference as trace 'I") at high orbit inclination
angle. ........................................................................................................................................ 13
Figure 17. Energy window contribution to TID from protons for the micro-satellite structures
A to H and unprotected solid-state device................................................................................ 14
Figure 18. Energy window contribution to TID from electrons for the micro-satellite
structures A to H and unprotected solid-state device ............................................................... 15
Figure 19. Relative TID values inside the micro-satellite housings in electron dose dominated
environment. The ratio is taken with respect to benchmarked housing 407............................. 16
Figure 20. Relative TID values inside the micro-satellite housings in proton dose dominated
environment. The ratio is taken with respect to benchmarked housing 407............................. 16
Figure 21. Ratio of TID values of micro-sat structure E with respect to structure F for protons
and electrons. Shielding effect of adding 4 mils of tantalum (structure F) is evident in case of
electrons.................................................................................................................................... 19
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List of tables
Table 1. List of model structures, shielding thickness and micro-satellite model bus mass ...... 6
Table 2. Proton to Electron Dose ratio for micro-satellite structures A to H and exposed solidstate device at 30 degrees inclination orbits ............................................................................ 10
Table 3. Proton to Electron Dose ratio for micro-satellite structures A to H and exposed solidstate device at 60 degrees inclination orbits ............................................................................ 10
Table 4. Proton to Electron Dose ratio for micro-satellite structures A to H and exposed solidstate device at 85 degrees inclination orbits ............................................................................ 11
Table 5. Total ionizing dose (TID) in Rad(Si) Y-1 from trapped radiation for micro-satellite
structures A to H and exposed solid-state device (“I”) in orbits having 30 degrees inclination.
Mass of the model structures is compared relative to structure "A” ....................................... 18
Table 6. Total ionizing dose (TID) in Rad(Si) Y-1 from trapped radiation for micro-satellite
structures A to H and exposed solid-state device (“I”) in orbits having 60 degrees inclination.
Mass of the model structures is compared relative to structure "A” ....................................... 18
Table 7. Total ionizing dose (TID) in Rad(Si) Y-1 from trapped radiation for micro-satellite
structures A to H and exposed solid-state device (“I”) in orbits having 85 degrees inclination.
Mass of the model structures is compared relative to structure "A” ....................................... 19
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1.
INTRODUCTION
Parts of the Earth’s magnetosphere that are capable of trapping ionizing radiation form the socalled “Earth’s Radiation Belts”. Occupying predominantly the inner portion of the
magnetosphere, the trapped radiation display spatial and temporal variation in their spectra.
The inner part of the radiation belts (L< 2.5), and the outer part (3 < L < 12) contain trapped
electrons and protons, the proton belt being predominantly confined to smaller L values, L
being the McIlwain magnetic shell parameter. In the inner part of the radiation belts, the
particle population density is more stable, while the outer part is variable and the particle
density responds readily to solar wind activity. The void region, located between the two parts
of the belts, gets filled during large magnetospheric activity.
Satellites located in Earth’s orbit will be required to operate in these radiation belts and in the
process will be exposed to trapped ionizing radiation, galactic cosmic rays and solar flare
radiation. Since the radiation is dependent on altitude and latitude, satellite orbits such as for
example LEO to GEO transfer orbit will pass through various regimes of radiation belts
involving different electron and proton spectra. For some of these orbits, and at different
orbital points, the radiation could be either proton or electron dose dominated. The integrated
space radiation environment per orbit will depend on a number of parameters such as the
spectrum of the radiation and the time of exposure to that spectrum and the local temporal
modulation by the magnetospheric activity. This is specifically true for elliptic orbits where
the velocity of the spacecraft will change at various points of the trajectory. In addition,
geomagnetic shielding and Earth shadowing will modulate exposure of the satellite to
radiation originating outside the Earth’s magnetosphere, specifically solar flare radiation and
galactic cosmic rays.
The total ionizing dose (TID) environment of the mission, to a very large extent, will be
affected by the design of the micro-satellite bus. Material composition of exterior walls,
relative location inside the satellite, location of solar cells with respect to the main microsatellite body, presence/absence of equipment housing, cable harness locations and many
other structural features will affect TID value. In order that the solid-state devices operating
on board a micro-satellite can meet the TID requirement of the mission, as a precursory step
at the design stage of the mission it is necessary to carry out mission dose estimates.
In this work, we will examine how the structure of the micro-satellite bus, relative location
inside the bus, material selection and orbit parameters affect the TID environment inside a
satellite. A schematic picture of the model micro-satellite is shown in Figure 1[1]. The
electronic housings have been numbered, as shown in Figure 1; the numbers also reflect the
ID number that the specific housing unit has inside the input file of the Monte Carlo radiation
transport simulation code used in this work. The objective of this work is to compare
shielding effectiveness of several micro-satellite shielding configurations for selected orbits
against trapped radiation and solar flare protons.
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1
417
407
427
307
317
Figure 1. Schematic picture of the interior of the model micro-satellite, showing the locations of the
electronic housings.
2.
MICRO-SATELLITE STRUCTURE SCHEMES
Several micro-satellite bus structures have been modeled to ascertain their shielding capability
against TID at selected orbits. The TID mitigation effort is often compared against the weight
budget of spacecraft, these two being opposing factors. Various shielding schemes are
employable for sensitive components protection, each placing different amounts of shielding
between the radiation environment and the radiation sensitive electronic solid-state device.
The solid-state device model is made of a 10mil thick silicon layer enclosed into a 100mil
thick molded epoxy package. Eight micro-satellite structures have been studied; the structures
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and materials are shown in Figures 2, 3, 4, 5 and 6. The dose deposition is calculated in the
model solid-state device, one in each of the 5 electronic housings. Throughout most of this
work, the results will refer to TID in the solid-state device located in the electronic equipment
housing numbered 407. The shielding effectiveness of other electronic housings is explored
later.
Structures A and B (Figure 2) can be termed as conventional buses with aluminum electronic
housings and aluminum structural panels. Electronic housings in structure B are 2mm thick;
all other structures (A, C, D, E and F) have the electronic housings 1mm thick. The supporting
shelves are also 1mm thick, made of aluminum. The structures C to H have the body panels
made of honeycomb mesh, covered on the outside and inside with thin layers of material, such
as aluminum or PEEK. The body panels of structure C are made of 8mm thick aluminum
honeycomb mesh covered on both sides with 0.1mm aluminum sheeting. Structure D has the
honeycomb portion of the body panels made of PEEK, a carbon composite material, which is
covered with thin, 0.1mm layer of aluminum sheeting on both sides (see Figure 3). Satellite
structure E is like structure D, except the honeycomb sheeting is made also from PEEK
material. Structure F is like structure E, however, on the inside of the body panels, there is a
0.1mm thick layer of tantalum, a high Z material (Figure 4). Structure G has no support
shelves because the electronic housings are attached directly to the spacecraft structural body
panels (Figure 5). Structure H has no electronic housings to house sensitive devices (Figure
6). Instead, the rigid electronic housing and electronic boards are replaced with flex-boards
and flex-cables that are attached directly to the structural body panels of the micro-satellite[2].
An electronic solid-state device would in this configuration be attached directly to the flexboard as shown in Figure 6. Table 1 provides a summary of shielding thickness and the mass
of the micro-satellite structures. The mass reflects only the mass of the supporting structure
and excludes all subsystems.
Structure "A"
Structural
Structral
pannels
panels
1mm Al
Electronic
housings 1mm Al
Structure "B"
1mm Al
2mm Al
Figure 2. Traditional aluminum micro-satellite bus structures schemes. Structure “B” models “spot”
shielding by increasing the electronic housing thickness.
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Structure "C"
8mm Al Honeycomb
0.1mm Al
0.1mm Al
Structure "D"
8mm PEEK Honeycomb
0.1mm Al
0.1mm Al
Structural
Structral
panels
pannels
Electronic
housings
1mm Al
1mm Al
Figure 3. The solid aluminum panel is replaced with a honeycomb aluminum panel in “C” and with a
carbon composite PEEK honeycomb in “D”. Both honeycomb panels are covered with 0.1mm aluminum
sheeting inside and outside. The electronic housings are 1mm thick solid aluminum.
Structure "E"
8mm PEEK Honeycomb
0.5mm PEEK
0.5mm PEEK
Structural
Structral
panels
pannels
Electronic
housings
Structure "F"
8mm PEEK Honeycomb
0.5mm PEEK
0.5mm PEEK
exterior
0.1mm Tantalum
1mm Al
1mm Al
Figure 4. Carbon composite PEEK has replaced aluminum material in structural panels; aluminum
electronic housings, however, remain. A thin coating (0.1mm) of high Z value material tantalum is added
to the interior surface in “F”.
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Structure "G"
8mm PEEK Honeycomb
0.5mm PEEK
0.5mm PEEK
Structural
Structral
panels
pannels
exterior
0.1mm Tantalum
Electronic
equipment
housing to
Electronic
equipment
housing attached
attaced
to the structural pannels
the structural
panels
Electronic
housings
1mm Al
Figure 5. Electronic housings are attached to the structural panels. The panels have the same
composition as in structure “F”.
Structure “H”H
Structure
Satellite Body Structural Panel
Satellite body pannel
0.1mm Tantalum / 0.5mm PEEK / 8mm PEEK Honeycomb / 0.5mm PEEK
0.1mm Tantalum /0.5mm PEEK / 8mm PEEK Honeycomb /0.5mm PEEK
Kapton flex-board
exterior
Silicon detector
encapsulated in Epoxy
Figure 6. Model representation of the flex-board structure panel. Aluminum housings have been
eliminated.
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Table 1. List of model structures, shielding thickness and micro-satellite model bus mass
Mass of the Model Supporting
Model Structure
Effective Shielding Thickness
Structure
0.813 g cm-2
6850g
A
B
1.084 g cm-2
C
-2
+ Al Honeycomb
4530g
-2
D
0.596 g cm
+ PEEK Honeycomb
5090g
E
0.662 g cm-2 + PEEK Honeycomb
5500g
F
3.
0.596 g cm
8790g
-2
+ PEEK Honeycomb
7500g
-2
0.822 g cm
G
0.551 g cm
+ PEEK Honeycomb
5990g
H
0.340 g cm-2 + PEEK Honeycomb
4050g
SHIELDING EFFECTIVENESS
Shielding effectiveness of bus structures is examined at 5 elliptical orbits with a common
perigee of 600km and with apogees of 1100km, 1500km, 3000km, 20000km and 36000km
(LEO to GEO transfer orbit) respectively. Three inclination angles, low (30 degrees),
medium (60 degrees) and high (85 degrees) have been used. The particle data were obtained
from the SPENVIS system, ESA’s space environment software package, available to run on
the World Wide Web. The radiation transport simulation code used in this study was the 3D
Monte Carlo code MCNPX/LAHET. The code is a general-purpose time-dependent transport
code for neutrons, photons, and electrons in combination with the LAHET module; it also
calculates transport and interaction of nucleons, pions, muons, light ions, and antinucleons.
The micro-satellite bus structures provide radiation protection of varying degree for the
sensitive electronic devices located on board, as shown in Figures 7 to 9 for trapped electron
radiation and in Figures 10 to 12 for the trapped protons. The contributions to TID are shown
separately for the purpose of ascertaining the relative shielding effectiveness of these
structures in both the electron and proton environments. This is useful as, for example, many
military missions require fission electron dose analysis for electronics on board. For the sake
of completeness, the TID for the exposed solid-state device is also shown. It is evident that in
the trapped electron environment, the least protection is provided by the bus structure type H;
however even this bus reduces the total annual electron dose by about an order of magnitude
from what a fully exposed device would receive. Bus structure B, with 1mm thick body
panels and 2mm thick aluminum electronic circuit housings, and structure F, with light polycarbon honeycomb body panels and a 0.1mm (4 mils) tantalum layer, provide the best
protection. A thin, 4 mils tantalum layer is virtually as much effective shield as an extra 1mm
of aluminum added to the aluminum housing for electrons. The other bus structures (A, C, D,
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E and G) are comparable to each other in shielding effectiveness. Utilization of honeycomb
body panels provides just as effective shielding against electrons as solid aluminum panels if
one compares the results of structure A against structures C, D, E and G. Electronic housings
can be used as an effective means of “spot” shielding by comparing results between structure
G and H, where structure H is effectively like G but without the electronic housing or
structure B and A, where structure B has thicker electronic housing.
In the proton environment, the shielding effectiveness differences of the structures are less
pronounced; only structure H and the unprotected device show distinct TID values and even
this difference decreases with increasing orbit of inclination. This is shown in Figures 10, 11
and 12, showing results of TID from trapped protons into the micro-satellite structures A to H.
Trace “I” is again the dose that a totally unprotected device would receive in these orbits.
The electron dose peaks in all cases at the orbit with the apogee of 20000km and becomes
lowest for LEO (1100km apogee). This is what one would expect from the distribution of the
trapped electron population inside the inner magnetosphere, where the trapped electron
population is predominantly located at higher magnetic L shell values. The proton dose peaks
at lower orbits than the electron dose, again based on similar arguments that the trapped
proton population is located predominantly in the inner portion of the radiation belts.
Specifically, for the considered orbit examples, the orbit with apogee of 3000km has the
proton dose peak for all three inclination angles. The dominance of either proton or electron
TIDs inside the spacecraft is summarized in Tables 2, 3 and 4 as the ratio of proton/electron
dose. The data show both orbit dependency (vertical columns) and spacecraft structure
dependency, shown horizontally.
.
Figure 7. Calculated annual trapped electrons TID values into micro-satellite structures A to H and
unprotected solid-state device, trace ” I”, for LEO, MEO and LEO-GEO transfer orbits with inclination
angle of 30 degrees
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Figure 8. Calculated annual trapped electrons TID values into micro-satellite structures A to H and
unprotected solid-state device, trace “I”, for LEO, MEO and LEO-GEO transfer orbits with inclination
angle of 60 degrees.
Figure 9. Calculated annual trapped electrons TID values into micro-satellite structures A to H and
unprotected solid-state device, trace “I”, for LEO, MEO and LEO-GEO transfer orbits with inclination
angle of 85 degrees.
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Figure 10. Calculated trapped protons annual TID values into micro-satellite structures A to H and
unprotected solid-state device, trace “I”, for LEO, MEO and LEO-GEO transfer orbits with inclination
angle of 30 degrees.
Figure 11. Calculated trapped protons annual TID values into micro-satellite structures A to H and
unprotected solid-state device, trace ‘I”, for LEO, MEO and LEO-GEO transfer orbits with inclination
angle of 60 degrees.
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Figure 12. Calculated trapped protons annual TID values into micro-satellite structures A to H and
unprotected solid-state device, trace “I”, for LEO, MEO and LEO-GEO transfer orbits with inclination
angle of 85 degrees.
Table 2. Proton to Electron Dose ratio for micro-satellite structures A to H and exposed solid-state
device at 30 degrees inclination orbits
ORBIT
APOGEE
A
B
C
D
E
F
G
H
I
1100km
1500km
3000km
20000km
20.99
57.65
19.23
21.76
26.21
67.98
27.53
9.826
1.010
13.29
36.15
12.20
13.78
16.56
42.55
17.46
6.445
0.715
5.552
15.36
5.136
5.768
6.897
17.68
7.483
3.305
0.693
0.026
0.061
0.024
0.027
0.031
0.066
0.033
0.018
0.007
36000km
0.036
0.091
0.033
0.037
0.044
0.097
0.047
0.022
0.005
Table 3. Proton to Electron Dose ratio for micro-satellite structures A to H and exposed solid-state
device at 60 degrees inclination orbits
ORBIT
APOGEE
A
B
C
D
E
F
G
H
I
1100km
1500km
3000km
20000km
3.126
8.062
2.902
3.253
3.871
8.451
3.760
1.455
0.231
5.457
14.27
5.051
5.672
6.766
15.52
6.779
2.632
0.386
4.640
12.25
4.305
4.813
5.718
14.07
6.261
2.889
0.665
36000km
0.001
0.003
0.001
0.001
0.001
0.003
0.002
0.001
0.0001
0.009
0.027
0.009
0.010
0.012
0.028
0.012
0.004
0.0005
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Table 4. Proton to Electron Dose ratio for micro-satellite structures A to H and exposed solid-state
device at 85 degrees inclination orbits
ORBIT
APOGEE
A
B
C
D
E
F
G
H
I
1100km
1500km
3000km
20000km
3.395
8.813
3.150
3.533
4.209
9.269
4.097
1.576
0.245
2.876
7.465
2.668
2.993
3.566
7.817
3.465
1.332
0.212
4.673
12.35
4.335
4.847
5.759
14.20
6.317
2.914
0.667
36000km
0.001
0.002
0.001
0.001
0.001
0.002
0.001
0.0003
0.00007
0.001
0.023
0.007
0.008
0.001
0.024
0.010
0.003
0.00032
The amount of relative shielding individual structures can provide is presented in Figure 13,
showing TID ratios at five locations inside the spacecraft; the locations are numbered after the
individual electronic housings as shown in Figure 1 (Further discussion to relative shielding
by individual housings is given later). The TID ratio is taken with respect to a fully exposed
solid-state device to the space environment at the specific orbit. Two orbit examples are
given; one is a high inclination angle, high apogee orbit (lower cluster of curves) and the other
is a low inclination angle, low apogee orbit (upper cluster of curves). As shown, the microsatellite bus structures can cut the TID values (in comparison to exposed solid-state device)
down to between 7% and 15% in low inclination, low altitude orbit environment and down to
between less then 2% and 6% in the high inclination angle, high altitude orbit environment.
Figure 13. Relative shielding effectiveness of the micro-satellite model structures A to F for two specific
orbit scenarios. The ratio is with respect to the fully exposed solid-state device.
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4.
SOLAR FLARE EFFECT
The JPL-91 solar flare model, available from SPENVIS, ESA’s space environment model,
was used to determine the solar proton fluence into the micro-satellite structures for the five
elliptic orbits used in this work. The mission was assumed to be five years long, geomagnetic
shielding was taken into account and the magnetosphere was considered to be stormy. The
stormy magnetospheric condition will provide greater geomagnetic shielding at the low
energy end of the proton spectrum up to about 50 MeV. Figures 14, 15, and 16 show the
annual TID contribution into the micro-satellite test structures, specifically into the polymer
encapsulated solid-state device located in the electronic housing numbered 407 (see Figure 1).
Figure 14. Calculated solar proton annual TID values into micro-satellite model structures A to H and
exposed solid-state device (shown for reference as trace “I”). No contribution to TID is observed at low
altitude orbits and low angle of inclination.
Again, for comparison purposes, the fully exposed solid-state device is also shown. For the
externally (external to magnetosphere) originating radiation, such as solar flare protons,
geomagnetic shielding is very effective at low inclination angles and low apogee orbits. At
the 30 degrees orbital inclination angle, effective screening for the selected orbits is well
beyond the 3000km apogee. At the 20000km apogee orbit, only solar protons with energy
greater then 60MeV deposit some dose into the satellite. The proton fluence is, however, low
at these energies and therefore contribution to TID is also low. The LEO-to GEO orbit
receives contributions to TID from all the energies, although this is smaller than it would be
for GEO because the micro-satellite becomes geomagnetically shielded in the vicinity of the
orbit perigee. At higher inclination angles, geomagnetic shielding is less effective and more
low energy solar protons contribute to TID. Shielding against solar flare protons, in these
examples, the traditional aluminum bus structure B provides marginally the best protection,
other structures are similar as shown in Figures 14, 15 and 16. The multifunction PEEK
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polycarbon/flex-board structure (H) offers the lowest protection against solar flare protons.
The unprotected, fully exposed solid-state device is also shown.
Figure 15. Calculated solar proton annual TID values into micro-satellite structures A to H and
unprotected solid-state device (shown for reference as trace “I”) at medium orbit inclination angle.
Figure 16. Calculated solar proton annual TID values into micro-satellite structures A to H and
unprotected solid-state device (shown for reference as trace “I”) at high orbit inclination angle.
DRDC Ottawa TM 2003-017
13
5.
ENERGY WINDOW CONTRIBUTION TO TID
All structures display a threshold energy for protons and electrons at which there is a large
jump in dose deposition. Below this energy, very little dose deposition occurs into any of the
solid-state devices located inside the micro-satellite bus structure. This threshold energy is bus
structure dependent; for the bus structures under consideration, the protons threshold is
between 15MeV and 25MeV while for electrons the threshold energy lies between 1MeV and
1.5MeV. The results are shown in Figures 17 and 18, showing the contribution to TID as a
function of proton and electron energy.
It is also evident that for the high-energy protons and electrons the structure configuration
becomes less important as for all the micro-satellite bus structures considered, the TID as a
function of bus structure design and material converges into a single value. This indicates that
shielding against high-energy particles, especially protons, which contribute to Single Event
Effects (SEE), is very difficult and is not very feasible. However, the good news is that the
population of trapped high energy protons is about 3 to 4 orders of magnitude less than the
population at the low-energy end of the spectrum. For both electrons and protons, again
configuration H requires the least energetic particles, the threshold energy being about 1 MeV
for electrons and 15 MeV for protons. For bus type B, the threshold energy was found to be
highest at 1.5MeV and 25MeV for electrons and protons respectively.
Figure 17. Energy window contribution to TID from protons for the micro-satellite structures A to H and
unprotected solid-state device.
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Figure 18. Energy window contribution to TID from electrons for the micro-satellite structures A to H
and unprotected solid-state device.
6.
LOCATION SHIELDING EFFECTVENESS
The shielding capability of the bus structure in other electronic housings (other than 407, see
Figure 1) inside the micro-satellite structure type A was examined. Analyses were carried out
for two types of orbits, namely, one with low angle of inclination and low altitude, and the
other with high angle of inclination and high altitude. The results are shown in Figures 19 and
20, presented as the ratio of TIDs taken with respect to the TID in the electronic housing
number 407. This electronic housing (407), as mentioned, has been used throughout this
work as the reference location.
Larger variation in TID values from location to location inside the micro-satellite is observed
in the high inclination and high altitude case, i.e. in the electron-dose-dominated environment
than in the proton dose environment case (low inclination, low altitude orbit) as evident if one
compares results in Figure 19 and 20. This indicates that shielding by location would be more
feasible in the electron dose dominated environment then in the proton environment. For the
orbits, without adding any additional shielding, the TID values can vary by up to 100% as
seen in Figure 19. The smaller variation at orbits with low altitude and low inclination angles
points to larger difficulty of the bus structures to shield against protons.
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15
Figure 19. Relative TID values inside the micro-satellite housings in high altitude and high inclination
angle orbit. The ratio is taken with respect to benchmarked housing 407.
Figure 20. Relative TID values inside the micro-satellite housings in low altitude and low inclination
angle orbit. The ratio is taken with respect to benchmarked housing 407.
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7.
DISCUSSION
Traditional micro-satellite bus structures having structural panels, shelves and equipment
housings made of aluminum and bus structures made of lighter materials such as polycarbon have been compared for their space radiation shielding effectiveness. Structure B,
with effective 3mm minimum aluminum shielding was the best in protection against the
space environment ionizing radiation, due to the additional 1mm of aluminum shielding
added to the equipment housing. Other structures with identical interiors provided good
comparison in terms of the micro-satellite envelope performance. These results are
summarized in Tables 5, 6 and 7. Multi-layered structure F with poly-carbon honeycomb
structural panels and a 4 mils tantalum layer provided the second-best protection against the
space environment ionizing radiation. By removing the 4 mils thick tantalum layer,
structure F becomes structure E, and the shielding protection at high altitude orbits becomes
reduced by a factor of between 2 and 3 (Figure 21 and Tables 2, 3 and 4), while shielding
protection show only a small deterioration at the low altitude orbits. The high Z material
(tantalum) is located in the particle path between low Z materials of the outside panel, i.e.
poly-carbon PEEK structural panel, and the aluminum housing. Tantalum provides selfattenuation for bremsstrahlung X-rays while the poly-carbon structural panels reduce
electron fluence via the inelastic scattering process [3].
Micro-satellite structures C and D are identical with the exception that the outer body panels
of structure D, specifically the honeycomb portion, is made of composite carbon material
PEEK (density 1.2g cm-3) compared to the honeycomb made of traditional aluminum
(density = 2.7g cm-3) in structure C. In the high altitude orbit environment, the D structure
performs slightly better; in the low altitude orbit environment, the two are about the same.
In terms of the weight budget, about 0.5kg reduction can be realized with structure C; even
though PEEK is only about as half as dense aluminum, thicker PEEK material making the
honeycomb panel (0.1mm versus 0.5mm) accounts for the difference. The presence of other
bus structures such as shelves, reduce the TID values by up to 60%, as was observed
comparing results between structure F and structure G, however the weight budget increases
by about 1.5kg. Although the presence of electronic housings increased the weight budget
of the model micro-satellite by about 2 kg, the total TID was reduced by up to a factor of 5
in high altitude, high inclination orbits. In the low altitude orbit environment, a reduction
was also observed, but was less than a factor of 2. The most dramatic change in TID occurs
when both shelves and electronic housings are removed; the effect is shown in Tables 5, 6,
and 7 by comparing results between structure F and structure H. Again, the change is most
visible in the high altitude orbit environment.
Protecting a sensitive electronic component by selecting a more shielded location inside a
micro-satellite can be done relatively well in the high altitude orbit environment but it
becomes a much less effective technique in the low altitude and low angle of inclination
orbit environments. For example, up to a 100% change in TID can occur from location to
location in structure A, by selecting a different location to house sensitive electronic device
as evident for example between TID results inside housings #317 and #427. Again, much
less change in protection by location can be accomplished when the micro-satellite is
located in an orbit with low angle of inclination and low altitude.
DRDC Ottawa TM 2003-017
17
Orbit parameters determine either electron or proton dose domination, but micro-satellite
structure also plays a role as to whether TID into the electronic components of the microsatellite is electron or proton dominated. There is a rather large (in some cases a factor of 7)
variation from structure to structure for the same orbit, as can be seen, which can be tied to
the previously made point that layered structures can be designed for electron shielding.
The orbital dependence is a function of spectrum change from orbit to orbit. Every microsatellite structure has also a different capability for shielding out particles up to a certain
particle energy. Threshold energies were determined at which a large, 4 to 5 orders of
magnitude, jump in contribution to TID occurs. These values were determined for all of the
micro-satellite structures and for both proton and electron radiation. Above this energy, the
contribution to TID becomes much less energy and micro-satellite design dependent for
both types of particles.
Table 5. Total ionizing dose (TID) in Rad(Si) Y-1 from trapped radiation for micro-satellite
structures A to H and exposed solid-state device (“I”) in orbits having 30 degrees
inclination. Mass of the model structures is compared relative to structure “A”.
Structure
Mass Ratio
ORBIT
(Ref. “A”)
1100km 1500km
3000km
20000km 36000km
9010
24060
13430
1890
A
1.00
496
3100
8610
10280
1630
437
B
1.28
9770
25810
13670
1900
498
C
0.66
8610
22860
13170
1870
491
D
0.74
7060
18980
12490
1820
480
E
0.80
3220
8980
11150
1720
462
F
1.10
7840
21370
14190
1950
510
G
0.87
28520
70270
25590
2780
687
H
0.59
314200
514030
107960
8020
1600
I
Table 6. Total ionizing dose (TID) in Rad(Si) Y-1 from trapped radiation for micro-satellite
structures A to H and exposed solid-state device (“I”) in orbits having 60 degrees
inclination. Mass of the model structures is compared relative to structure “A”.
Structure
Mass Ratio
ORBIT
(Ref. “A”)
1100km 1500km
3000km
20000km 36000km
3260
19500
7490
1340
A
1.00
402
1020
6560
5590
1080
309
B
1.28
3580
21010
7640
1350
409
C
0.66
3130
18510
7330
1310
394
D
0.74
2520
15230
6920
1260
374
E
0.80
1040
6840
6080
1150
327
F
1.10
2740
17110
7930
1360
402
G
0.87
11300
58910
14900
2160
690
H
0.59
166400
471770
66560
8040
2960
I
-
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DRDC Ottawa TM 2003-017
Table 7. Total ionizing dose (TID) in Rad(Si) Y-1 from trapped radiation for micro-satellite
structures A to H and exposed solid-state device (“I”) in orbits having 85 degrees
inclination. Mass of the model structures is compared relative to structure “A”.
ORBIT
Structure
Mass ratio
(Ref. “A”) 1100km
1500km
3000km
20000km 36000km
1930
16510
6440
448
A
1.00
357
595
5450
4810
340
277
B
1.28
2130
17820
6570
457
363
C
0.66
1860
15680
6310
439
350
D
0.74
1490
12850
5950
415
332
E
0.80
599
5690
5230
359
293
F
1.10
1600
14420
6820
447
358
G
0.87
6840
50850
12840
785
605
H
0.59
113500
421630
57550
3540
2550
I
-
Figure 21. Ratio of TID values of micro-satellite structure E with respect to structure F for
protons and electrons. Shielding effect of adding 0.16g cm-2 of Tantalum (structure F) is
well evident in case of electrons.
DRDC Ottawa TM 2003-017
19
8. SUMMARY
Radiation transport analysis into several micro-satellite bus structures was presented. The
results indicate that it is possible to design a bus structure, optimal to operate in a specific
orbit space environment. In the high-altitude orbit environment, multi-layered structure made
of layers of low Z material and high Z material provides very effective protection against the
total ionizing dose. In the low-altitude orbit environment, the presence of such layering is not
necessary; structures made only of low Z materials are just as effective in shielding.
9. REFERENCES
1. “QuickSat” Space Technologies Micro-satellite Platform Development Project,
CSA/DND Working Group Presentation, November 5, 1999 Meeting
2. B.D. Spieth, K.S. Quasim, R.N. Pottman and D.A. Russell, “Shielding Electronics
Behind Composite Structures”, IEEE Transactions on Nuclear Science, Vol. 45,
No. 6, December 1998.
3. W.C. Fan, C.R. Drumm, S.B. Roeske and G.J. Scrivner, “Shielding Consideration
for Satellite Microelectronics”, IEEE Transactions on Nuclear Science, Vol. 43,
No. 6, December 1996
20
DRDC Ottawa TM 2003-017
List of
symbols/abbreviations/acronyms/initialisms
DND
TID
LEO
MEO
GEO
Department of National Defence
Total Ionizing Dose
Low Earth Orbit
Mid-altitude Earth Orbit
Geostatinary Earth Orbit
DRDC Ottawa TM 2003-017
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22
DRDC Ottawa TM 2003-017
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Evaluation of Electronic Shielding in Micro-satellites(U)
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Varga, L and Horvath, E
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Abstract
This report investigates the shielding capabilities of several micro-satellite model structures and shielding materials in
protecting internal spacecraft electronics from the ionizing radiation of the space environment. The calculations have
been carried out with 3D Monte Carlo radiation transport code. The results indicate that the largest reduction of total
ionizing dose (TID) is observed with traditional aluminium spacecraft structures, although, structures made with lighter
poly-carbon materials with added thin layer of high-Z material can provide comparable radiation protection in addition
to much desired spacecraft mass reduction.
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LEO, SEU, PEEK, Space Environment, Tantalum, Trapped Protons, Trapped Electrons, Solar Protons, Total
Ionizing Dose
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